The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases that flow downstream through a high pressure turbine nozzle which directs the flow into a row of high pressure turbine rotor blades. The blades extract energy from the gases for powering the compressor, and a low pressure turbine is disposed downstream therefrom for extracting additional energy which typically powers a fan for producing propulsion thrust to power an aircraft in flight.
The high pressure turbine nozzle receives the highest temperature combustion gases directly from the combustor and is specifically configured for withstanding those gases for a useful service life. The nozzle is an annular structure but is segmented into arcuate segments to accommodate the substantial expansion and contraction of the components thereof due to the hot operating environment. Each segment includes arcuate outer and inner bands supporting a pair of hollow stator vanes which receive a portion of pressurized air bled from the compressor for cooling the nozzle segments during operation.
The two bands define the radially outer and inner flowpath surfaces between which the combustion gases are confined during operation. The bands are separated from each other by corresponding axial splitlines which are suitably sealed with typical spline seals therebetween.
The nozzle vanes have a crescent profile with substantial curvature or camber between the leading and trailing edges thereof, with a generally concave pressure side and a generally convex opposite suction side along which the combustion gases flow during operation. The suction side of one vane is circumferentially spaced from the pressure side of an adjacent vane to define a flow channel therebetween for the combustion gases. The combustion gases enter these flow channels in a general axial downstream direction and are redirected at an oblique angle from the outlet of the channels defined between adjacent vane trailing edges.
Accordingly, the individual streamlines of the combustion gases flow generally parallel to each other between the nozzle vanes, but vary in curvature to correspond with the different velocities thereof as effected by the suction and pressure sides of adjacent vanes.
The band splitlines are straight and oriented obliquely in the bands between the corresponding arcuate profiles of the adjacent vane suction and pressure sides. Accordingly, the combustion gases typically cross the splitline twice during their passage between the vanes as they curve between the suction and pressure sides and flow axially aft along the splitlines.
The bands are circumferentially continuous between each pair of vanes in each nozzle segment and enjoy maximum aerodynamic efficiency. However, the splitlines between the vanes of adjacent nozzle segments provide a local discontinuity in the bands which can affect aerodynamic efficiency.
The band flow surfaces are designed to be substantially flush with each other at the splitlines, but due to normal manufacturing tolerances and stack-up of those tolerances during assembly of the nozzle components, differences in radial elevation of the adjoining bands randomly occur with corresponding steps in the flow path surfaces. If the step faces forwardly opposite to the direction of the combustion gases, they introduce a local obstacle to the smooth flow of those gases which both reduces aerodynamic efficiency of the nozzle and locally heats the exposed edges leading to oxidation thereof over time. Oxidation of the exposed splitline edges reduces the useful life of the nozzle segments and requires earlier replacement thereof than would be otherwise required.
Since the combustion gases typically cross the oblique splitlines twice as they pass through the nozzle channels, the undesirable flow-obstructing steps may occur at either the forward portion of the bands or the aft portion of the bands or may vary therebetween in a transition zone therebetween. Since a typical aircraft gas turbine engine operates over different power levels from idle to maximum power, the configuration of the streamlines through the nozzle correspondingly varies.
Accordingly, a downstream facing step during some operation of the engine may change to an upstream facing step as the configuration of the streamlines changes. Undesirable oxidation of the upstream facing edge remains a practical problem due to real-life manufacturing tolerances and the inability to exactly dimension the nozzle components. Nozzle durability is therefore affected by the exposed splitline edges which shortens the useful life of the nozzle in practice.
It is, therefore, desired to provide an improved turbine nozzle including an improved splitline configuration for enhancing durability and useful life of the nozzle.
A turbine nozzle includes segments of outer and inner bands supporting corresponding vane pairs. The bands adjoin each other at corresponding ends along splitlines, with each band having a forward land, an opposite aft land, and a middle land extending therebetween. The forward lands have a nominal aft-facing step, the aft lands have a nominal forward-facing step and the middle lands are nominally flush.